This invention relates to spacecraft attitude control, and more particularly to three-axis spacecraft attitude control using a polar star sensor in conjunction with an Earth sensor.
Artificial satellites or spacecraft are in widespread use for various purposes. For some purposes, as for example for communication purposes, a satellite may be required to direct an instrument toward particular locations on the surface of a heavenly body about which it orbits, such as Earth.
For communications between spaced-apart locations on the surface of the Earth, or for broadcasting purposes, the geosynchronous satellite provides certain advantages. The geosynchronous satellite occupies an equatorial orbit, and may be either spin-stabilized or three-axis stabilized. In order to keep an antenna or antennas pointed at particular locations, the spin-stabilized satellite must include a de-spun platform on which the Earth-pointing antennas or sensors are mounted. Spin stabilized spacecraft require large momentum biases to maintain tight pointing.
A three-axis stabilized spacecraft uses a system of torquers for applying torques to the spacecraft in relation to roll, pitch and yaw axes for maintaining its attitude under the control of one or more attitude sensors. Torquing to maintain orientation may be accomplished by magnetic torquers, by chemical thrusters, by momentum or reaction wheels or any combination of the three.
The attitude of the spacecraft may be determined by any of a number of systems. One attitude sensing system, used on commercial communications satellites, requires an Earth horizon sensor together with a sun sensor for attitude control. The Earth sensors provide pitch and roll information, while the sun sensor provides information at certain times of day which allows for updating of the yaw information. Yaw must be estimated during those intervals in which a sun sensor measurement is not available. These sensors, together with the quarter-orbit interchange of roll and yaw axes, provide sufficient information for stabilization of a spacecraft under ordinary conditions. The horizon sensors, however, may not provide precise information because of atmospheric affects in obscuring the horizon, and the yaw update information may not provide sufficient control for precise pointing. For less precise pointing requirements on momentum bias spacecraft only the horizon sensors are needed and the gyroscopic coupling of the momentum wheel may be used to maintain the yaw attitude.
In another attitude control system, gyroscopes may be mounted within the spacecraft to provide angular rate information. Since gyroscopes are subject to errors due at least to bearing friction, they must be periodically updated. The updates may be provided by Earth horizon sensors and sun sensors, but these are subject to errors as described above. The gyros may be updated by a system of star sensors. Such a system of star sensors observes portions of the sky and matches the observed scene with a memorized star map. Such an arrangement may provide precise control, but is extremely expensive because the star sensor must respond to stars over a substantial range of magnitudes in order to make a meaningful comparison with the memorized chart. The requirement for observation of stars over a substantial range of magnitudes in turn requires cooling of the image sensing devices in order to detect the fainter stars, and also requires linearity of response or at least a response which is well characterized so that the relative magnitudes of the star may be determined for the comparison. The cost of such a star sensing system is prohibitive for a commercial communications satellite. Gyroscopes suitable for long life such as might be required for a 10-year commercial communications satellite application are extremely expensive. The cost of such gyros is also prohibitive for commercial communications satellites.